Inner diffuser case struts for a combustor of a gas turbine engine

ABSTRACT

An inner diffuser case for use in a combustor of a gas turbine engine includes a multiple of struts across an annular flow path between an outer shroud and an inner shroud, at least one of the multiple of struts is an open strut.

Applicant hereby claims priority to PCT Patent Application No.PCT/US2013/062145 filed Sep. 27, 2013, which claims priority to U.S.Patent Application No. 61/707,016 filed Sep. 28, 2012, the disclosuresof which are herein incorporated by reference.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to a combustor section therefor.

Gas turbine engines, such as those that power modern commercial andmilitary aircraft, generally include a compressor to pressurize anairflow, a combustor for burning a hydrocarbon fuel in the presence ofthe pressurized air, and a turbine to extract energy from the resultantcombustion gases. The compressed air from the compressor section passesbetween stator vanes, then through a diffuser section. The diffuser hasan expanding cross sectional area in the direction of the airflow todecrease the velocity and increase the static pressure of the air. Thisprepares the air for entry into a combustion section at low velocity topermit proper mixing with fuel.

In gas turbine engines of the twin spool type, leakage of high pressurecompressor discharge air through a labyrinth seal between a highpressure compressor rear rotor shaft and the diffuser section typicallyestablishes the thrust loads on component parts such as thrust bearings.

This air is also used downstream in the engine for cooling purposes,particularly, the bearing compartment and the high pressure turbine. Airthat is cooler than the environment surrounding the bearing compartment,is bled from the mid-stage compressor either through an anti-vortex tubedirecting the compressor bleed air radially inward to the engine's shaftwhere it flows axially to the environment surrounding the mid-framebearing compartment, or externally from the mid high pressure compressorcase to the bearing compartment, then used to supply cooling to portionsof the high pressure turbine rotor.

SUMMARY

An inner diffuser case for use in a combustor of a gas turbine engineaccording to one disclosed non-limiting embodiment of the presentdisclosure includes an outer shroud, an inner shroud, an annular flowpath defined by and between the outer and inner shrouds and a multipleof struts extending across the annular flow path wherein at least one ofsaid multiple of struts is an open strut.

In a further embodiment of the foregoing embodiment, the multiple ofstruts include a multiple of open struts. In the alternative oradditionally thereto, in the foregoing embodiment the multiple of openstruts communicate a secondary flow air. In the alternative oradditionally thereto, in the foregoing embodiment the secondary flow airprovides a thrust balance.

A combustor section of a gas turbine engine according to anotherdisclosed non-limiting embodiment of the present disclosure includes aninner diffuser case with at least one open strut.

In a further embodiment of the foregoing embodiment, the inner diffusercase defines an annular flow path. In the alternative or additionallythereto, in the foregoing embodiment the at least one open strut extendsacross said annular flow path.

In a further embodiment of any of the foregoing embodiments, thecombustor section includes a forward assembly of a combustor immediatelydownstream of said inner diffuser case.

In a further embodiment of any of the foregoing embodiments, thecombustor section includes an outer diffuser case mounted to said innerdiffuser case.

A method of communicating secondary flow air in a combustor section of agas turbine engine according to another disclosed non-limitingembodiment of the present disclosure includes communicating a secondaryflow air through a strut across an annular flow path of an innerdiffuser case.

In a further embodiment of the foregoing embodiment, the method includesproviding a thrust balance with the secondary flow air.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is a partial longitudinal schematic sectional view of a combustorsection according to one non-limiting embodiment that may be used withthe gas turbine engine shown in FIG. 1;

FIG. 3 is a partial longitudinal schematic sectional view of an innerdiffuser case;

FIG. 4 is a front view of the inner diffuser case;

FIG. 5 is a sectional view of an open strut of the inner diffuser casetaken along line 5-5 in FIG. 4;

FIG. 6 is a sectional view of a solid strut of the inner diffuser casetaken along line 6-6 in FIG. 4; and

FIG. 7 is a schematic view of the open strut of the inner diffuser case.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengines such as a turbojets, turboshafts, and three-spool (plus fan)turbofans wherein an intermediate spool includes an intermediatepressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”)and a High Pressure Compressor (“HPC”), and an intermediate pressureturbine (“IPT”) between the high pressure turbine (“HPT”) and the Lowpressure Turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 (“LPC”) and a lowpressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42directly or through a geared architecture 48 to drive the fan 42 at alower speed than the low spool 30. An exemplary reduction transmissionis an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A whichis collinear with their longitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with thefuel and burned in the combustor 56, then expanded over the HPT 54 andthe LPT 46. The turbines 54, 46 rotationally drive the respective lowspool 30 and high spool 32 in response to the expansion. The main engineshafts 40, 50 are supported at a plurality of points by bearingstructures 38 within the static structure 36. It should be understoodthat various bearing structures 38 at various locations mayalternatively or additionally be provided.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 48can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3, and in another example is greater thanabout 2.5:1. The geared turbofan enables operation of the low spool 30at higher speeds which can increase the operational efficiency of thelow pressure compressor 44 and low pressure turbine 46 and renderincreased pressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine 46 is pressuremeasured prior to the inlet of the low pressure turbine 46 as related tothe pressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 44, and the low pressure turbine 46has a pressure ratio that is greater than about five (5:1). It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentdisclosure is applicable to other gas turbine engines including directdrive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path B due to the high bypass ratio. The fan section 22 ofthe gas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the gas turbine engine 20 at its best fuelconsumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“T”/518.7^(0.5)) in which “T” represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, the combustor 56 generally includes an outercombustor wall 60, an inner combustor wall 62 and a diffuser case module64. The outer combustor wall 60 and the inner combustor wall 62 arespaced apart such that a combustion chamber 66 is defined therebetween.The combustion chamber 66 is generally annular in shape. The outercombustor wall 60 is spaced radially inward from an outer diffuser case64-O of the diffuser case module 64 to define an annular outer plenum76. The inner combustor wall 62 is spaced radially outward from an innerdiffuser case 64-I of the diffuser case module 64 to define an annularinner plenum 78. It should be understood that although a particularcombustor is illustrated, other combustor types with various combustorliner arrangements will also benefit herefrom. It should be furtherunderstood that the disclosed cooling flow paths are but an illustratedembodiment and should not be limited only thereto.

The combustor walls 60, 62 contain the combustion products for directiontoward the turbine section 28. Each combustor wall 60, 62 generallyincludes a respective support shell 68, 70 which supports one or moreliners 72, 74 mounted to a hot side of the respective support shell 68,70. The liners 72, 74, often referred to as Impingement Film Float (IFF)wall panels define a generally rectilinear liner array which define theannular combustor chamber 66. Each of the liners 72, 74 may be generallyrectilinear and manufactured of, for example, a nickel based superalloy, ceramic or other temperature resistant material. In one disclosednon-limiting embodiment, the array includes a multiple of forward liners72A and a multiple of aft liners 72B that line the hot side of the outershell 68 and a multiple of forward liners 74A and a multiple of aftliners 74B that line the hot side of the inner shell 70.

The combustor 56 further includes a forward assembly 80 immediatelydownstream of the compressor section 24 to receive compressed airflowtherefrom. The forward assembly 80 generally includes an annular hood82, a bulkhead assembly 84, a multiple of fuel nozzles 86 (one shown)and a multiple of fuel nozzle guides 90 (one shown). Each of the fuelnozzle guides 90 is circumferentially aligned with one of the hood ports94 to project through the bulkhead assembly 84. Each bulkhead assembly84 includes a bulkhead support shell 96 secured to the combustor walls60, 62, and a multiple of circumferentially distributed bulkheadheatshields segments 98 secured to the bulkhead support shell 96 aroundthe central opening 92.

The annular hood 82 extends radially between, and is secured to, theforwardmost ends of the combustor walls 60, 62. The annular hood 82includes a multiple of circumferentially distributed hood ports 94 thataccommodate the respective fuel nozzle 86 and introduce air into theforward end of the combustion chamber 66 through a central opening 92.Each fuel nozzle 86 may be secured to the diffuser case module 64 andproject through one of the hood ports 94 and through the central opening92 within the respective fuel nozzle guide 90.

The forward assembly 80 introduces core combustion air into the forwardend of the combustion chamber 66 while the remainder enters the annularouter plenum 76 and the annular inner plenum 78. The multiple of fuelnozzles 86 and surrounding structure generate a blended fuel-air mixturethat supports combustion in the combustion chamber 66.

Opposite the forward assembly 80, the outer and inner support shells 68,70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in theHPT 54. In one disclosed non-limiting embodiment, thirty-two (32) NGVs54A are located immediately downstream of the combustor 56 as the firststatic vane structure upstream of a first turbine rotor in the turbinesection 28. The NGVs 54A are static engine components which direct coreairflow combustion gases onto the turbine blades of the first turbinerotor in the turbine section 28 to facilitate the conversion of pressureenergy into kinetic energy. The core airflow combustion gases are alsoaccelerated by the NGVs 54A because of their convergent shape and aretypically given a “spin” or a “swirl” in the direction of turbine rotorrotation. The turbine rotor blades absorb this energy to drive theturbine rotor at high speed.

With reference to FIG. 3, the inner diffuser case 64-I defines anannular flow path 100 (FIG. 4) for compressed airflow C from theupstream HPC 52. The annular flow path 100 includes a multiple of struts102 which extend in a radial direction between an outer shroud 104 andan inner shroud 106

With reference to FIG. 4, the multiple of struts 102 of the innerdiffuser case 64-I include a multiple of open struts 102A (FIG. 5) and amultiple of solid struts 102B (FIG. 6). It should be appreciated thatalthough a particular number of open struts 102A and solid struts 102Bare illustrated, any number will benefit herefrom.

With reference to FIG. 7, the multiple of open struts 102A provides ametered flow path with passages 108 for secondary flow air illustratedschematically by arrows S). This secondary flow air S is typicallysourced from the compressor section 24 and may be utilized, for example,as the reaction force to the axial forces on the high spool 32 (FIG. 1)such that the bearing structures 38 (see FIG. 1) on which the outershaft 50 is supported has an acceptable level of axial load or thrustbalance. Generally, aerodynamic and pressure forces exert an axial loadon the outer shaft 50 through the compressor section 24 and the turbinesection 28. While the compressor section 24 driven by the turbinesection 28 can compensate to some degree for a net axially downstreamforce in the turbine section 28, a rotor thrust bearing, as designatedgenerally by the numeral 38, is normally required to fully absorb theaxial thrust forces. The axial loads absorbed by the bearing structures38 is thereby assisted or supplemented by differential pressuresgenerated with a metered flow of the secondary flow air S on each sideof the compressor section 24 and/or the turbine section 28.Consequently, the maximum thrust force which is to be absorbed by thethrust bearings is reduced.

In order to achieve this axial thrust balance, the secondary flow air Sis metered by the relationship of, for example, the number of openstruts 102A (FIG. 5) with respect to the multiple of solid struts 102Bas well as the size of the passages 108 to provide a specific flow areato obtain the desired pressure at the desired locations. That is, thetotal cross sectional area of the passages 108 of the open struts 102Amay be utilized to control the high compressor module rear side/aftpressure as desired.

The inner diffuser case 64-I strut arrangement incorporates passages 108for secondary flow air S which locate the metering function directlyinto the part detail casting which reduces cost, increases the functionsof the inner diffuser case 64-I and provides an efficient secondary airflow path.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. An inner diffuser case for use in a combustor ofa gas turbine engine comprising: an outer shroud; an inner shroud; anannular flow path defined by and between the outer and inner shrouds;and a multiple of struts extending across the annular flow path, themultiple of struts including a multiple of open struts and a multiple ofsolid struts; and wherein a first of the multiple of open struts isarranged circumferentially between and adjacent to an adjacent pair ofthe multiple of solid struts.
 2. An inner diffuser case for use in acombustor of a gas turbine engine comprising: an outer shroud; an innershroud; an annular flow path defined by and between the outer and innershrouds; and a multiple of struts extending across the annular flowpath, the multiple of struts including a multiple of open struts and amultiple of solid struts; and wherein a first of the multiple of solidstruts is arranged circumferentially between and adjacent to a first ofthe multiple of open struts and a second of the multiple of solidstruts; and the second of the multiple of solid struts is arrangedcircumferentially between and adjacent to the first of the multiple ofsolid struts and a second of the multiple of open struts.